Composite structures for aerospace vehicles, and associated systems and methods

ABSTRACT

Composite structures for space vehicles and other aerospace vehicles are disclosed herein. In one embodiment of the disclosure, a space vehicle pressure vessel includes composite panels or sidewalls with septumized core. In another embodiment, a splice joint for joining septumized composite panels in space vehicle structures is disclosed herein.

CROSS-REFERENCE TO RELATED APPLICATION(S) INCORPORATED BY REFERENCE

The present application is a continuation application of U.S. patentapplication Ser. No. 12/885,346, filed Sep. 17, 2010, entitled COMPOSITESTRUCTURES FOR AEROSPACE VEHICLES. AND ASSOCIATED SYSTEMS AND METHODS,which claims priority to U.S. Provisional Patent Application No.61/244,006, filed Sep. 18, 2009, and entitled “COMPOSITE STRUCTURES FORAEROSPACE VEHICLES, AND ASSOCIATED SYSTEMS AND METHODS,” and U.S.Provisional Patent Application No. 61/244,008, filed Sep. 18, 2009, andentitled “COMPOSITE STRUCTURES FOR AEROSPACE VEHICLES. AND ASSOCIATEDSYSTEMS AND METHODS.” each of which is incorporated herein in itsentirety by reference.

TECHNICAL FIELD

The present disclosure is directed generally to composite structuresand, more particularly, to composite structures for space vehicles andassociated systems and methods.

BACKGROUND

Rocket powered launch vehicles carry humans and other payloads intospace. For example, rockets took the first humans to the moon andreturned them safely home. Rockets also launch satellites and unmannedspace probes, and carry supplies and personnel to the internationalspace station. Despite the rapid advances in manned and unmanned spaceflight, however, delivering astronauts, satellites, and other payloadsto space continues to be an expensive proposition.

Although NASA's space shuttle is largely reusable, reconditioning thereusable components is a costly and time consuming process that requiresextensive ground based infrastructure. Moreover, the additional shuttlesystems required for reentry and landing reduce the payload capabilityof the Shuttle.

Pressure vessels on conventional spacecraft (e.g., payload capsules,crew capsules, tanks, etc.) are typically made of aluminum, stainlesssteel, titanium, and/or other metals which can be relatively heavy,expensive, or both. Because of weight restrictions, capsules aretypically made of a single layer of material, which may provide limitedstructural redundancy in the event of damage. Aspects of the presentdisclosure are directed to addressing these challenges.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1A is a schematic side view of a space vehicle structure configuredin accordance with an embodiment of the disclosure, and FIG. 1B is apartially exploded schematic side view of the space vehicle structure ofFIG. 1A.

FIG. 2 is an enlarged cross-sectional view of a composite sandwich panelconfigured in accordance with an embodiment of the disclosure, and FIG.2A is an enlarged cross-sectional view of a composite sandwich panelconfigured in accordance with another embodiment of the disclosure.

FIGS. 3A-3F are cross-sectional views illustrating various stages in amethod of forming a structural splice joint in a composite sandwichstructure in accordance with an embodiment of the disclosure.

FIGS. 4A-4F are a series of plan views illustrating various stages in amethod of bonding a doubler between two composite panels in accordancewith another embodiment of the disclosure.

FIGS. 5A-5C are a series of cross-sectional views illustrating variousstages in a method of reinforcing composite sandwich panels with pottingcompound to accept a bolt or other suitable fastener in accordance witha further embodiment of the disclosure.

DETAILED DESCRIPTION

The present disclosure is directed generally to composite structures,such as septumized composite structures for space vehicles. Severaldetails describing structures and processes that are well-known andoften associated with composite materials and composite partmanufacturing techniques are not set forth in the following descriptionto avoid unnecessarily obscuring embodiments of the disclosure.Moreover, although the following disclosure sets forth severalembodiments, several other embodiments can have differentconfigurations, arrangements, and/or components than those described inthis section. In particular, other embodiments may have additionalelements, and/or may lack one or more of the elements described belowwith reference to FIGS. 1A-5C.

FIG. 1A is a schematic diagram illustrating a composite structureconfigured in accordance with an embodiment of the disclosure, and FIG.1B is a partially expanded schematic diagram of the composite structure.Referring first to FIG. 1A, in the illustrated embodiment the compositestructure is a space vehicle structure 100, such as a suborbital ororbital capsule or other pressure vessel. More specifically, in theillustrated embodiment the space vehicle structure 100 is a capsulehaving a crew cabin 110. A forward section, such as a nose section 102can be attached to a forward end portion of the crew cabin 110, and anaft bulkhead or aft section 104 can be similarly attached toward an aftend of the crew cabin 110.

Referring next to FIG. 1B, in the illustrated embodiment the crew cabin110 is a hollow pressure vessel having a forward bulkhead or cap 114sealably spliced to a forward end of a body or sidewall section 112, andan aft bulkhead or cap 116 sealably spliced to the aft end of thesidewall section 112. As described in greater detail below, the sidewallsection 112, the forward cap 114 and/or the aft cap 116 can beconstructed of a septumized composite sandwich that provides increaseddamage tolerance, reduced weight, and structural redundancy.

In the illustrated embodiment, the sidewall section 112 can be formedfrom two or more relatively large barrel or annular ring sections whichare spliced together in accordance with various embodiments of thedisclosure described in detail below. For example, the sidewall section112 can be formed from a first ring section 112 a that is joined to asecond ring section 112 b along a circumferential joint 113. Althoughthe sidewall section 112 of the illustrated embodiment has a conicalshape, in other embodiments, other composite pressure vessels configuredin accordance with the present disclosure can have other shapes,including cylindrical shapes, curved or ogive shapes, etc. Moreover,although the composite structure of the illustrated embodiment is aspace vehicle structure 100, the methods, systems and/or structuresdisclosed herein can also be used to make and use other types ofcomposite structures, such as other types of pressure vessels,hypersonic vehicles, aircraft, surface ships and underwater vehicles.Accordingly, the present disclosure is not limited to space vehiclestructures.

FIG. 2 is an enlarged cross-sectional view of a portion of the sidewallsection 112 configured in accordance with an embodiment of thedisclosure. In the illustrated embodiment, the sidewall section 112includes a first face sheet 220 spaced apart from a second face sheet222, and an intermediate layer or sheet 224 positioned between the firstface sheet 220 and the second face sheet 222 to form a septum. Thesidewall section 112 further includes a first core portion 230adhesively bonded between the first face sheet 220 and the intermediatesheet 224, and a second core portion 232 adhesively bonded between theintermediate sheet 224 and the second face sheet 222.

In the illustrated embodiment, the outer or first face sheet 220, theinner or second face sheet 222, and the intermediate sheet 224 can becomprised of a plurality of laminated plies of fiber reinforced resinmaterial, such as fiber fabric reinforced epoxy resin. In oneembodiment, for example, the laminated plies can include graphite/epoxy,pre-preg, and woven fabric such as Toray T700SC-12 K-50C/#2510. In otherembodiments, at least the first face sheet 220 can include a layer ofmaterial with a relatively high ballistic modulus or resistance todamage from ballistic impact, such as a layer of Kevlar® material(para-aramid synthetic fiber), and/or a layer of electrically conductivematerial. Moreover, any one of the first face sheet 220, the second facesheet 222, and/or the intermediate sheet 224 can include one or morelayers of low porosity material to at least reduce pressure leakage fromthe pressure vessel. In other embodiments, however, other types oflaminated material layers can be used to form the face sheets 220 and222, and the intermediate sheet 224.

In another aspect of this embodiment, the first core portion 230 and thesecond core portion 232 can be comprised of a honeycomb core, such as aphenolic or aluminum honeycomb core that is perforated to facilitateventing. For example, the first core portion 230 and the second coreportion 232 can be comprised of Hexcel Aluminum Standard CoreCRLLL-1/8-5056-002 perforated; and/or Hexcel Aluminum Flex CoreCRLLL-5052/F40-0037S 5.7 perforated. In the illustrated embodiment, thefirst core portion 230 and the second core portion 232 can be about 0.5inch thick, resulting in an overall panel width of from about 0.75 inchto about 1.5 inches, or about 0.9 inch to about 1.2 inches, or about 1inch. In other embodiments, the first core portion 230 and the secondcore portion 232 can be comprised of other suitable core materials,including other suitable honeycomb materials, and closed and/or opencell foam materials, and can have other thicknesses.

The sidewall section 112 can be manufactured in one embodiment by firstlaying up one or more plies of composite fabric (e.g., pre-preggraphite/epoxy fabric) on a suitable lay up mandrel to form the inner orsecond face sheet 222, and then applying a layer of adhesive, such asEA-9696 epoxy film adhesive from Hysol (not shown) to the outer surfaceof the second face sheet 222. Next, the second core portion 232 can bepositioned on the second face sheet 222, and another layer of adhesivecan be applied to the outer surface of the second core portion 232. Theone or more plies of material making up the intermediate sheet 224 canthen be positioned on the second core portion 232, and another layer ofadhesive can be applied to the outer surface of the intermediate sheet224. The first core portion 230 can then be positioned on theintermediate sheet 224, and another layer of adhesive can be positionedon the outer surface of the first core portion 230. The one or moreplies making up the first face sheet 220 can then be positioned on theouter surface of the first core portion 230, and the assembly can bevacuum-bagged and debulked for curing using suitable methods and systemsknown in the art.

The foregoing construction of the sidewall section 112 creates astructural member which can be described as a first composite sandwichlayer 226 and a second composite sandwich layer 228. The redundancyprovided by the dual sandwich layers can increase damage resistance. Forexample, if the outer face sheet 220 and adjacent core portion 230 aredamaged from, for example, a micro-meteor strike, the inner sandwichlayer 228 can carry the design limit loads without failure. Although thesidewall section 112 of the illustrated embodiment includes a singleseptum (e.g., the septum formed by the intermediate sheet 224), those ofordinary skill in the art will appreciate that other sidewall sectionsand composite structures configured in accordance with the presentdisclosure can include more than one septum without departing from thespirit or scope of the present disclosure. For example, in otherembodiments the methods and systems disclosed herein can be used makecomposite panels and other structures having two, three or more septumsdepending on the needs of the particular application. FIG. 2A, forexample, is an enlarged cross-sectional view of a portion of a sidewallsection 112 a having second and third intermediate sheets 224 a and 224b, respectively, and third and fourth core portions 234 and 236,respectively. The second and third intermediate sheets 224 a and 224 bare spaced apart from the first intermediate sheet 224 and disposedbetween the first face sheet 220 and the second face sheet 222. Thethird core portion 234 is disposed between the first intermediate sheet224 and the second intermediate sheet 224 a, and the fourth core portion236 is disposed between the second intermediate sheet 224 a and thethird intermediate sheet 224 b.

FIGS. 3A-3F are a series of cross-sectional views illustrating variousstages in a method of forming a composite panel splice joint inaccordance with an embodiment of the disclosure. By way of example, inone embodiment the splice joint described with reference to FIGS. 3A-3Fand suitable variations thereof can be used to form the splice joint 113between the first composite section 112 a and the second compositesection 112 b shown in FIG. 1B. Referring first to FIG. 3A, an edgeportion of the first outer sandwich layer 226 a includes a first rampsurface 334 a formed from the first face sheet 220 a. Similarly, an edgeportion of the second inner sandwich layer 228 b includes a second rampsurface 334 b made up of the second face sheet 222 b. Prior to bondingthe first composite section 112 a to the second composite section 112 b,a first faying surface portion 340 a of the first inner sandwich layer228 a is grit blasted with, for example, an aluminum oxide or othersuitable medium to clean the surface 340 a and prepare it for adhesivebonding. A second faying surface portion 340 b of the second outersandwich layer 226 b is similarly grit blasted.

Referring next to FIG. 3B, adhesive 352, such as an epoxy paste adhesivelike Hysol EA-9394, or other suitable adhesives known in the art, isapplied to the faying surfaces of the first inner sandwich layer 228 aand the second outer sandwich layer 226 b. The edge portion of thesecond outer sandwich layer 226 b is then positioned over the edgeportion of the first inner sandwich layer 228 a as shown in FIG. 3B. Inthe illustrated embodiment, a gap of from about 0.02 inch to about 0.06inch, such as about 0.04 inch, can exist between the faying surfaces ofthe first composite section 112 a and the second composite section 112 bfor bonding. In addition, a first edge 350 a of the first inner sandwichlayer 228 a can be spaced apart from a second edge 350 b at the bottomof the ramp surface 334 b by a distance from about 0.15 inch to about0.4 inch, or about 0.25 inch. In other embodiments, the faying surfacesof the composite sections 112 and/or the distance between the respectiveedges can vary depending on the particular structural application ormanufacturing method employed.

In one particular embodiment, a spacer, such as a filament 360 (e.g., around filament shown in end view and FIG. 3B) can be positioned betweenthe faying surfaces 340 a and 340 b to control the gap therebetweenwhile the adhesive 352 is curing. In other embodiments, other systemsand methods can be used to control the relative position of thecomposite sections 112 during the bonding process.

Referring next to FIG. 3C, after the composite sections 112 have beenbonded together, a suitable potting compound, such as L-306 pottingcompound provided by JD Lincoln Company of 851, West 18th Street, CostaMesa, Calif. 92627, can be used to fill the gaps between the opposingedge portions of the composite sandwiched layers 226 and 228. In oneembodiment, the L-306 potting compound is a lightweight,room-temperature curing, “edge filler” compound. In other embodiments,other suitable potting compounds, adhesives and/or other fillermaterials can be used to fill the gaps between the composite sandwichedlayer edge portions as shown in FIG. 3C.

Referring next to FIG. 3D, after the potting compound 370 has cured, theexterior surfaces of the potting compound can be sanded smooth, and thena first surface area 372 a on the first face sheet 220 can be gritblasted with a suitable media, such as aluminum oxide, and a secondsurface area 372 b on the second face sheet 222 can be similarly gritblasted. In the illustrated embodiment, the prepared surfaces 372 canextend for about two inches on either side of the compound 370. Gritblasting the surfaces in this manner can help prepare and clean thesurface for proper adhesion of subsequent adhesive applications.

Referring next to FIG. 3E, one or more dry fabric plies 380 a, such asgraphite fabric plies, can be applied over the adjacent edge portions ofthe first and second outer composite sandwich layers 226 a, 226 b, andone or more dry fabric plies 380 b can be similarly applied over theadjacent edge portions of the first and second inner composite sandwichlayers 228 a, 228 b. In one embodiment, the dry fabric plies 380 caninclude dry graphite fabric overwraps such as material type BMS9-8 Type1 Class II Style 3K-70-PW. In other embodiments, other types of drygraphite fabric and/or fabric pre-impregnated with resin can be used tooverwrap the joint and provide a structural splice.

In the embodiment where dry graphite fabric is used, a suitable resin,such as a suitable epoxy laminating resin such as Hysol EA-9396, isapplied to the dry graphite fabric to infuse the fabric with the resin.After the fabric and resin have been applied as shown in FIG. 3E, afirst vacuum bag 390 a can be installed over the first doubler plies 380a, and a second vacuum bag 390 b can be installed over the seconddoubler plies 380 b. The vacuum bags 390 can then be evacuated and/orde-bulked to compress the splice joint 113 between the doubler plies 380for curing of the resin. In other embodiments, the splice joint 113and/or the doubler plies 380 can be cured in a suitable autoclave. Asshown in FIG. 3E, in one embodiment the doublers 380 can extend fromabout one inch to about three inches, or about 1.5 inches over theadjacent edge portions of the respective composite panel sections 112.In other embodiments, the doublers 380 can have other dimensions.

Referring next to FIG. 3F, this figure is a cross-sectional view of asplice joint 310 that joins two septumized composite sandwich panelstogether using a method that is at least generally similar to the methoddescribed above in reference to FIGS. 3A-3E.

FIGS. 4A-4F are a series of plan views illustrating various stages inthe method for bonding a doubler splice 480 to adjacent edge portions oftwo or more composite sandwich panels 412 in accordance with anembodiment of the disclosure. The doubler bonding process describedherein can be used to bond a doubler to virtually any type of surface ormember, not just composite panels. For example, the methods describedherein can be used to bond a suitable doubler to, e.g., metal panels,metal/composite panels, organic material panels, etc.

Referring first to FIG. 4A, as this view illustrates the edges of thefirst and second panel sections 412 a and 412 b are positioned adjacentto each other in generally planar alignment. In one embodiment, thebonding process begins by determining the periphery of a vacuum bagaround the doubler 480 for application of bagging edge tape (e.g., zincchromate) or other suitable sealing material or device 484. The bag seal484 should bridge over a portion of the doubler 480 in the vicinity ofthe gap between the adjacent structural panels 412 as shown in FIG. 4A.Next, the doubler 480 can be covered with a protective material, such astape (not shown), on the non-bonded surface. The doubler 480 and theprotective material can then be drilled as shown to produce the holes482. In addition to the foregoing steps, the portion of the compositepanels 412 that will not be bonded to the doubler 480 can also becovered by a protective material, such as tape, where a bond is notdesired. Additional steps can include positioning a bagging film 490over the doubler 480 and attaching a suitable evacuating device to avacuum port 496. These and other steps are described in more detailbelow with reference to FIGS. 4B-4F.

Referring next to FIG. 4B, breather material, such as rolled-up breathermaterial 492 can be positioned on the doubler 480 in between theadjacent rows of holes 482 (FIG. 4A) on the non-bonded side of thedoubler 480. The rolled breather material 492 can be taped in place.Next, a suitable adhesive, such as an epoxy film adhesive, can beuniformly applied to the faying surfaces between the doubler 480 and thecomposite panels 412. The doubler 480 can then be positioned in place onthe composite panels 412 and secured with tape if necessary. Adhesive494 may flow into the holes 482 (FIG. 4A) as shown in FIG. 4C. Vacuumdevice supports 496 are also shown in FIG. 4C.

As shown in FIG. 4D, an additional portion of breather material 493,such as breather material folded in the form of a blanket or pad, canalso be positioned over the rolled up breather material 492 (FIG. 4B).Referring to FIGS. 4D and 4E together, the bagging film 490 can then bepositioned over the breather material 493 and sealed against the edgetape (e.g., zinc-chromate edge tape) 484. A vacuum port 497 is thenoperably and sealably attached to the bagging film 490 and the volumeunderneath the film is evacuated, or at least partially evacuated, for apredetermined amount of time required to let the adhesive 494 cure. Avacuum gauge 498 can be suitably attached to the bagging film 490 asshown on FIGS. 4E and 4F to monitor the pressure on the doubler 480during the curing process. After the adhesive is cured, the bagging film490, sealing tape 484, and breather materials 492 and 493 can beremoved, and excess material can be cleaned from the bonded joint.

FIGS. 5A-5C are a series of cross sectional views illustrating variousstages in the method of applying potting compound to composite sandwichpanels 512 to enable the panels 512 to be fastened together with asuitable fastener 560 in accordance with an embodiment of thedisclosure. Referring first to FIG. 5A, in the illustrated embodiment apiece of core 530 is machined or otherwise hogged-out in a locationwhere it is desired to install a fastener. The opened area created bythe hog-out is then filled with potting compound 570 a as shown in FIG.5A. Next, a first face sheet 520 a and second face sheet 520 b can bebonded or otherwise laminated to the outer surfaces of the core 530using various suitable techniques known in the art.

Referring next to FIG. 5B, a second composite sandwich panel 512 b thatis at least generally similar in structure and function to the firstcomposite panel 512 a can be positioned adjacent to the first panel 512a so that the first portion of potting compound 570 a is adjacent to asecond portion of potting compound 570 b in the second composite panel512 b. A through-hole 540 can then be drilled or otherwise formedthrough the potting compounds 570, and a countersink or larger openingor hole 542 can then be formed from the backside of the second compositepanel 512 b.

Referring next to FIG. 5C, a nutplate 572, such as a domed or sealed nutplate, can be fastened and/or adhesively bonded into the recess createdby the second hole 542 with additional potting compound 574. As isknown, the nutplate 572 and similar fastening devices include a nut or asimilarly threadable engagement feature configured to receive malethreads on a bolt 560 or other suitable threaded fastener. In otherembodiments, potted-in type threaded inserts, such as an AEP Torloninsert (e.g., an AEP1035 floating nut, blind threaded, potted-in typeinsert from Marketing Masters, Inc. of 1871 NW Gilman Blvd., Issaquah,Wash. 98027) can be used in place of the nutplate 572. Once the nutplate572 has been bonded in place, the fastener 560 can be removed so thatthe composite panels 512 can be disassembled. Moreover, an overwraplayer or doubler ply 580 can be applied over the potting compound 574and bonded to the adjacent area of the second panel 512 b.

The potted structure illustrated in FIG. 5C enables the panels 512 to besubsequently joined together with only blind access to the outer side ofthe first composite panel 512 a. Moreover, the method of potting andfastening described above with reference to FIGS. 5A-5C can also beemployed for attaching equipment, structural members (e.g., a bracket562), and/or other objects to one or more composite panels.

From the foregoing, it will be appreciated that specific embodiments ofthe invention have been described herein for purposes of illustration,but that various modifications may be made without deviating from thespirit and scope of the various embodiments of the invention. Further,while various advantages associated with certain embodiments of theinvention have been described above in the context of those embodiments,other embodiments may also exhibit such advantages, and not allembodiments need necessarily exhibit such advantages to fall within thescope of the invention. Accordingly, the invention is not limited,except as by the appended claims.

We claim:
 1. A space vehicle configured to contain at least one of crewor payload, the space vehicle comprising: a septumized compositesandwich panel that forms a damage resistant outer wall of the spacevehicle, wherein the septumized composite sandwich panel comprises anouter face sheet that forms an exterior surface of the outer wall of thespace vehicle; an inner face sheet spaced apart from the outer facesheet; at least one intermediate sheet disposed between the outer andinner face sheets; a first core portion disposed between the outer facesheet and the intermediate sheet; and a second core portion disposedbetween the inner face sheet and the intermediate sheet, wherein theseptumized composite panel forms a redundant structure in which theintermediate sheet, the second core portion, and the inner face sheetare configured to perform without structural failure if the outer facesheet and the first core portion are damaged from object impact.
 2. Thespace vehicle of claim 1 wherein the outer and inner face sheets, theintermediate sheet, and the first and second core portions form apressure vessel.
 3. The space vehicle of claim 1 wherein the outer andinner face sheets, the intermediate sheet, and the first and second coreportions form a pressure vessel configured to maintain an internalpressure that is greater than the external pressure.
 4. The spacevehicle of claim 1 wherein the outer and inner face sheets, theintermediate sheet, and the first and second core portions form a spacevehicle capsule.
 5. The space vehicle of claim 1 wherein the outer facesheet, the inner face sheet, and the intermediate sheet are comprised offiber reinforced epoxy fabric, and wherein the first and second coreportions are comprised of a honeycomb core.
 6. The space vehicle ofclaim 1 wherein at least the outer face sheet includes a layer ofelectrically conductive material.
 7. The space vehicle of claim 1wherein at least the outer face sheet includes a layer of material witha high ballistic modulus.
 8. The space vehicle of claim 1 wherein atleast the outer face sheet includes a layer of para-aramid syntheticfiber material.
 9. The space vehicle of claim 1 wherein the spacevehicle is a pressure vessel, and wherein the septumized compositesandwich panel includes a low porosity sealing layer to at least reducepressure leakage from the pressure vessel.
 10. The space vehicle ofclaim 1, wherein the intermediate sheet is a first intermediate sheet,and wherein the septumized composite sandwich panel further comprises:at least a second intermediate sheet disposed between the outer andinner face sheets and spaced apart from the first intermediate sheet;and a third core portion disposed between the first intermediate sheetand the second intermediate sheet.
 11. The space vehicle of claim 1wherein the septumized composite sandwich panel further comprises: fouror more core portions disposed between the outer and inner face sheets;and three or more intermediate sheets alternatingly interposed betweeneach of the core portions.
 12. The space vehicle of claim 1 wherein theouter face sheet and the first core portion comprise a first sandwichpanel, wherein the inner face sheet and second core portion comprise asecond sandwich panel, and wherein the first sandwich panel is attachedto the second sandwich panel by a fastener extending through pottingcompound in the first and second sandwich panels.
 13. The space vehicleof claim 1 wherein the septumized composite sandwich panel forms acylindrical sidewall portion of the space vehicle, and wherein the spacevehicle is a pressure vessel.
 14. The space vehicle of claim 1 whereinthe septumized composite sandwich panel forms a cylindrical sidewallportion of the space vehicle, and wherein the space vehicle furthercomprises: a forward cap sealably attached to a forward end of thecylindrical sidewall portion; and an aft cap sealably attached to an aftend of the cylindrical sidewall portion, wherein the cylindricalsidewall portion, the forward cap, and the aft cap form a pressurevessel.
 15. The space vehicle of claim 1 wherein the septumizedcomposite sandwich panel forms a cylindrical sidewall portion of thespace vehicle, and wherein the space vehicle further comprises: aforward cap sealably attached to a forward end of the cylindricalsidewall portion, wherein the forward cap is formed from a septumizedcomposite sandwich structure; and an aft cap sealably attached to an aftend of the cylindrical sidewall portion, wherein the aft cap is formedfrom a septumized composite sandwich structure, and wherein thecylindrical sidewall portion, the forward cap, and the aft cap form apressure vessel.